1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to the cooling of airfoils in a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a compressor supplies compressed air to a combustor and burned with a fuel to produce a hot gas flow, which is then passed through a turbine to produce mechanical energy. The efficiency of the engine can be increased by passing a higher temperature flow through the turbine. The limiting factor is the temperature of the flow is the material properties used in the hot parts of the turbine. Typically, the rotor blades and stationary vanes of the first stage are exposed to the hottest gas flow. These parts are cooled by passing cooling air through complex passages formed within the airfoils. The engine efficiency can also be increased by using less cooling air flow through the cooled airfoils. The cooling air is usually bleed off air from the compressor. Use of bleed off air for cooling means less compressed air is available for combustion.
U.S. Pat. No. 5,702,232 issued to Moore on Dec. 30, 1997 entitled COOLED AIRFOILS FOR A GAS TURBINE ENGINE discloses an airfoil having a cooling supply channel formed by an inner wall of the airfoil (as represented in FIG. 1 of this application), and a plurality of radial feed passages positioned between the inner wall and the outer wall of the airfoil. Each feed passage is connected to the cooling supply passage by a re-supply hole, and each feed passage includes a film cooling hole connected to the airfoil outer surface. The Moore patent provides for near-wall cooling of the airfoil wall. However, this cooling construction, spanwise and chordwise cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. Also, a single pass radial channel flow is not the best method of utilizing cooling air, resulting is low convective cooling effectiveness.
U.S. Pat. No. 6,981,846 B2 issued to Liang on Jan. 3, 2006 entitled VORTEX COOLING OF TURBINE BLADES discloses an airfoil with a cooling supply passage formed by an inner wall of the airfoil (as represented in FIG. 2 of this application), and a plurality of radial extending vortex cooling chambers positioned between the inner wall and the outer wall of the airfoil. Three radial vortex chambers are connected in series, with the upstream-most chamber connected to the cooling supply channel and the downstream-most vortex chamber connected to a film cooling hole. The multi-vortex cell serves to generate a high coolant flow turbulence level and, hence, yields a very high internal convection cooling effectiveness in comparison to the single pass construction of the prior art. The Liang U.S. Pat. No. 6,981,846 B2 is incorporated herein by reference.
It is an object of the present invention to provide for a near-wall cooling for a turbine airfoil which will reduce the airfoil metal temperature and therefore reduce the cooling flow requirement and improve the turbine efficiency.